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Topic: Basic Rocket Science Q & A (Read 331781 times)

So, if the exhaust pressure is below the ambient pressure, you end up with a negative term in your thrust equation. i.e. (Pe - P0)*Ae < 0. And ergo, a reduction in thrust.

Yes: if the nozzle is over-expanded, then it is actually sucking the rocket backward to some extent. If this is hard to imagine, it's because our intuition about fluids is based entirely on subsonic flows.

Intrigued by the LANTR-concept, I tried to replicate the different Isps for different mixture ratios by using the formula Ve^2=2k/(k-1)*R*T/M, which I've found here and simplified a bit by removing the part with different pressures, as that factor would most likely be rather close to 1.

However, I always get results that are way off, e.g., if I go for pure hydrogen & T=2900, according to the 1st linkI should get an Isp of around 940, so a Ve of around 9000. But what I actually get is: Ve=sqrt(2*1.666/0.666*8.314*2900/1)=347.311m/s, which is off by a factor of 30(!).

Could anybody tell me what exactly I'm missing here?

If you're doing the calculation in SI units, than the molecular weight of atomic hydrogen is 0.001 kg/mol. At 2900 K, though, I would think most of the hydrogen would be molecular, so 0.002 kg/mol might be more accurate, and the specific heats would need to be adjusted too.

It's a relatively minor point, but ideal-gas specific heats may not be very accurate at high temperatures. As the temperature rises, molecules behave less and less like the rigid rotors, because vibrations and bending become significant.

An old but vague memory suddenly popped up. Which may very well be utterly wrong.

Somewhere amongst my hundred or so books on rocketry and spaceflight I'm sure there was a reference somewhere (probably dating to events between 1945 and 1970) describing cryogenic (oxygen I think) lines making a lot of scary noise during fueling operations. This might even be X-1 or X-15 territory.

If memory serves the noise was sort of continuous, not sudden bangs or anything and genuinely worrying to those not used to it.

But now I come to think about it I don't remember any launch coverage with audio containing anything much, other than the noise (before launch) of venting gases.

What are the reasons for using hydrogen in a first stage? Ever? This is a genuine question.

Ok, Shuttle did it, H-II does it, Ariane V does it. But all with (comparitively poor) Isp boosters.

Hydrogen is _great_ for an an upper stage, but surely a royal pain in the rear in the volumes required (and the necessary insulation for that immense volume) in the warm lower atmosphere for a (necessarily large) first stage.Never mind embrittlement of alloys et al.

Saturn used H2, but _only_ when high and fast in upper stages, Centaur ditto. No hydrogen low and slow.

Looking back, hydrolox in first (boosted) stages looks like a fad. Methalox _now_ looks like a fad, but perhaps a much better one, though that is to be proven.

I wonder as a mere software engineer if there ever was, or is, a good reason for a rocket engine to ever be burning hydrogen at close to sea level.

What are the reasons for using hydrogen in a first stage? Ever? This is a genuine question.

Ok, Shuttle did it, H-II does it, Ariane V does it. But all with (comparitively poor) Isp boosters.

Hydrogen is _great_ for an an upper stage, but surely a royal pain in the rear in the volumes required (and the necessary insulation for that immense volume) in the warm lower atmosphere for a (necessarily large) first stage.Never mind embrittlement of alloys et al.

Saturn used H2, but _only_ when high and fast in upper stages, Centaur ditto. No hydrogen low and slow.

Looking back, hydrolox in first (boosted) stages looks like a fad. Methalox _now_ looks like a fad, but perhaps a much better one, though that is to be proven.

I wonder as a mere software engineer if there ever was, or is, a good reason for a rocket engine to ever be burning hydrogen at close to sea level.

All three of your examples use large SRBs. For example, Ariane 5 gets 90% of its takeoff thrust from the solids. So, one way of looking at it is the SRBs are the first stage while the hydrolox booster is the second stage. No matter what you call it, when the SRBs are very large, a hydrolox booster makes sense.

What are the reasons for using hydrogen in a first stage? Ever? This is a genuine question.

Ok, Shuttle did it, H-II does it, Ariane V does it. But all with (comparitively poor) Isp boosters.

Hydrogen is _great_ for an an upper stage, but surely a royal pain in the rear in the volumes required (and the necessary insulation for that immense volume) in the warm lower atmosphere for a (necessarily large) first stage.Never mind embrittlement of alloys et al.

Saturn used H2, but _only_ when high and fast in upper stages, Centaur ditto. No hydrogen low and slow.

Looking back, hydrolox in first (boosted) stages looks like a fad. Methalox _now_ looks like a fad, but perhaps a much better one, though that is to be proven.

I wonder as a mere software engineer if there ever was, or is, a good reason for a rocket engine to ever be burning hydrogen at close to sea level.

All three of your examples use large SRBs. For example, Ariane 5 gets 90% of its takeoff thrust from the solids. So, one way of looking at it is the SRBs are the first stage while the hydrolox booster is the second stage. No matter what you call it, when the SRBs are very large, a hydrolox booster makes sense.

Another way to look at it is this - pretty much all rocket fuel choices involve a tradeoff between thrust, ISP, and propellant density (and thus stage dry mass).

With LH2, you get pretty good thrust, really good ISP, and not so good propellant density. The "opposite" is SRBs, where you get a whole lot of thrust, high propellant density, but comparatively low ISP. Kerolox is basically right in between the two when it comes to ISP, thrust, and prop density, with Methalox being similar to Kerolox but slightly less dense. It just so turns out that Kerolox strikes just a pretty optimal balance between thrust, ISP, and propellant density for 1st stages.

So...the end result is the overwhelming majority of rockets are either powered by kerolox 1st stages OR a combination of LH2 and SRBs, which, not coincidentally, yields an average propellant density and ISP between the two fuels as being about the same as Kerolox. The lone exception I can think of is the three core all LH2 Delta IV heavy, which has always seemed to me (visually at least) to really struggle off the pad.

What are the reasons for using hydrogen in a first stage? Ever? This is a genuine question.

Ok, Shuttle did it, H-II does it, Ariane V does it. But all with (comparitively poor) Isp boosters.

To this list we can add Energia. As RonM points out, in all of these cases hydrogen-fueled systems provide only a fraction of the take-off thrust, so were dealing not so much with first stages but with upper stages that happen to be lit on the ground.

As incoming says, Delta IV Medium and Delta IV Heavy do on the other hand have truly hydrogen-fueled first stages. My guess is that the sole reason is that at the time they went into development, the US had no modern hydrocarbon engines, having foolishly abandoned development of such after the F-1.

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Hydrogen is _great_ for an an upper stage, but surely a royal pain in the rear in the volumes required (and the necessary insulation for that immense volume) in the warm lower atmosphere for a (necessarily large) first stage.Never mind embrittlement of alloys et al.

I generally agree, though insulating a large hydrogen-fueled stage might in an important sense be easier than insulating a small one, because the large stage has a lower ratio of surface area to volume.

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Saturn used H2, but _only_ when high and fast in upper stages, Centaur ditto. No hydrogen low and slow.

If the parameters of the first stage are fixed, then a hydrogen-fueled upper stage maximizes payload. If the design of the first stage is a free parameter, however, hydrogen may not look as attractive. A hydrogen upper stage will give the smallest first stage, but it may not give the smallest, cheapest launch vehicle overall.

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Looking back, hydrolox in first (boosted) stages looks like a fad.

I wasn't there, but I'll bet many people were seduced by lox-hydrogen's high specific impulse and simply assumed it was the way to go, without careful analysis of hydrogen's drawbacks.

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Methalox _now_ looks like a fad, but perhaps a much better one, though that is to be proven.

I don't think methalox is a fad. Its major proponents -- Musk and Bezos -- are hard-headed businessmen who are probably more tightly focused on economic efficiency and less likely to be distracted by pure performance specs (e.g., specific impulse) than NASA or DoD.

Consider methane's advantages: methalox is quite a bit denser than hydrolox, and, though less dense than kerolox, it offers higher specific impulse. It's very cheap, and relatively easily made on Mars. In contrast to kerosene, methane is chemically very stable, making it suitable for staged-combustion cycles. Its viscosity is very low, reducing pressure losses. It can be stored at lox temperatures, allowing common-bulkhead tankage with insulation. Unlike kerosene, methane readily evaporates without leaving a residue, simplifying reuse of engines.

please, how can I calculate needed amount of oxygen in gaseous state versus liquid. If you need 1 unit of liquid oxygen, how many units of gaseous it will be? Under atmospheric pressure, or under 200 bar?

please, how can I calculate needed amount of oxygen in gaseous state versus liquid. If you need 1 unit of liquid oxygen, how many units of gaseous it will be? Under atmospheric pressure, or under 200 bar?

Thank you

pv=nrt is the equation you need I think. p is pressure, v is volume, n is number of moles of substance (set n1=n2 if trying to determine ratio of volumes at different v, p or t), r is the gas constant, and t is temp (in Kelvin). Gas constant is the same on both sides of the equation so thatíll cancel out so youíre left with p1v1/t1 = p2v2/t2. I think thatís all correct. Someone correct me if Iím off base. Itís been quite some time since Iíve delved into chemistry.

please, how can I calculate needed amount of oxygen in gaseous state versus liquid. If you need 1 unit of liquid oxygen, how many units of gaseous it will be? Under atmospheric pressure, or under 200 bar?

Thank you

Not sure what you mean by "unit."

If you mean mass, then it is easy, 1 kg of liquid oxygen is 1 kg of gaseous oxygen.

More likely you mean volume, in which case the answer is to determine the density and use that to determine how much mass you have in the one state (liquid or gas) and then use that mass and the density of the other state to determine the final volume.

For determining the volume of a gas, it depends on both temperature and pressure, so you need to use the ideal gas law that cppetrie mentioned. (density is n/ (V*M) where M is the molar mass of the substance)

For liquids, density is also a function of temperature, but at this point you are best off using a lookup table to determine the density at the conditions you care about. (Whether this temperature dependence is significant or not depends on the situation, it is generally a smaller effect than in a gas so it might not make much difference, but SpaceX cools their liquid oxygen extra cold to make it a bit denser and increase how much fuel fits in the rocket.)

please, how can I calculate needed amount of oxygen in gaseous state versus liquid. If you need 1 unit of liquid oxygen, how many units of gaseous it will be? Under atmospheric pressure, or under 200 bar?

Thank you

You have X liters of pure LOX.

1) Multiply by the density to find the mass of O2 present. 2) Divide by molecular mass of oxygen: 31.998g/mol 3) Multiply by molar volume of an ideal gas: 22.4L/mol (@1atm & 273K) 4) Adjust volume for pressure and temperature differences from STP. a) Set P1*V1/T1 = P2*V2/T2, and solve for V2.

The general thrust equation for rockets is: F=mdot*Ve + (Pe - Pa)*Ae where, F=thrust, mdot=mass flow rate, Ve=velocity of the exhaust at the nozzle exit, Pe=exhaust pressure at the nozzle exit, Pa=ambient pressure, and Ae=area of the nozzle exit.

When considering the Pa term during the low altitude segment of the flight, is that very strictly the general atmospheric pressure at the rocket's altitude? I believe the forward travel of the rocket and backward travel of the high velocity exhaust gasses creates a localized low pressure at the base of the rocket (this causes the plume recirculation seen on some launches, right?). Does this area's lowered pressure need to be taken into account for a higher accuracy calculation of the thrust?

Does anyone know how many rocket engine nozzles have been damaged in flight?

I know STS-93 experienced a frightening event just after liftoff that I understand involved the regen nozzle being punctured through, but are there any other known instances of such problems elsewhere?

Does anyone know how many rocket engine nozzles have been damaged in flight?

I know STS-93 experienced a frightening event just after liftoff that I understand involved the regen nozzle being punctured through, but are there any other known instances of such problems elsewhere?

Maybe the Merlin 1C that failed on an early Falcon 9 flight? (I seem to recall that involved implosion of the nozzle, but I'm not sure).

On CRS-1, a Merlin 1C engine had a fuel dome (a structure above the thrust chamber) rupture and lose pressure. Nothing to do with the nozzle.

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...When considering the Pa term during the low altitude segment of the flight, is that very strictly the general atmospheric pressure at the rocket's altitude? I believe the forward travel of the rocket and backward travel of the high velocity exhaust gasses creates a localized low pressure at the base of the rocket (this causes the plume recirculation seen on some launches, right?). Does this area's lowered pressure need to be taken into account for a higher accuracy calculation of the thrust?

If you want to talk about pressure variations along the rocket body, due to forward flight, then you are discussing terms that count as drag as well. The Pa term basically comes from the atmospheric pressure pushing back on the equivalent area at the front of the rocket, which obviously increases with velocity.

For accurate performance calculations, they would need to account for all of the various effects at the same time, including drag and variations in the thrust performance with altitude. I think your question can be rephrased as "does the thrust performance with altitude variation depend on forward velocity as well?" I think the answer to this is most likely yes, though I am not sure how significant it would be. Note that Pe and Ve vary as the exit flow shape changes, so the variation with altitude isn't simply due to the Pa term changing as one might think.